1. Field of the Invention
The present invention relates generally to spacecraft and, more particularly, to spacecraft solar arrays.
2. Description of the Related Art
Spacecraft typically carry solar cells as a primary energy source with rechargable batteries providing energy storage for eclipse operations. The solar cells are positioned on the spacecraft so that they are exposed to solar radiation.
On spinning spacecraft, solar cells are generally arranged about the outside of a spinning spacecraft body. Accordingly, only a fraction of the cells are exposed to solar radiation at any instant in time. On body-stabilized spacecraft, in contrast, solar cells are typically arranged in planar arrays and carried on solar wings which extend from opposite sides of a spacecraft body. Preferably, the solar wings rotate to keep them as orthogonal to the solar radiation as possible. Because the solar wings can be quite long in their deployed configuration, they generally are formed of a plurality of planar solar panels which are coupled together in an accordion arrangement so that they can be collapsed to a smaller stowed configuration for spacecraft launch.
The number of solar cells that must be carried by a spacecraft is a function of the anticipated spacecraft power drain and the efficiency of the solar cells. Although high-efficiency solar cells reduce the number of cells required by a specific spacecraft, they are quite expensive. Because weight and weight-related costs also increase with the number of solar cells, there is a considerable incentive to reduce the quantity of solar cells that a spacecraft must carry.
Accordingly, efforts have been extended to concentrate solar radiation upon solar cells with reflective surfaces that are positioned to reflect radiation upon the cells. Solar radiation that would otherwise have passed by a solar wing is thus redirected to be incident upon the solar cells. Energy conversion efficiency of this reflected radiation is typically less than that of direct radiation because of a lesser angle of incidence. The additional radiation also contributes to solar cell heating which decreases the radiation-to-electrical energy conversion efficiency. Nonetheless, the additional incident radiation facilitates a significant reduction in the number of spacecraft solar cells with consequent savings in spacecraft weight and cost.
A variety of reflector systems have been proposed. In an exemplary system of U.S. Pat. No. 4,282,394, reflector arms are carried on both inboard and outboard frames. Each of the reflector arms is formed of a plurality of hinged arm sections and each arm section of the inboard frame carries a reflective plastic sheet that is wound on a spring-biased roll. An end of each sheet is attached to a respective arm section on the outboard frame.
During deployment, an extensible shaft moves the outboard frame away from the inboard frame and each reflective sheet is unrolled to reflect solar radiation onto solar cells. Although this reflector system concentrates solar radiation, its complex structure (e.g., hinged arms, inboard and outboard frames and extensible shaft) significantly contributes to spacecraft weight and cost.
In a Naval Research Laboratory design, one-piece, thin-film reflectors are positioned on oppostie sides of a plurality of solar panels that are coupled together in an accordion arrangement. Edges of each thin-film reflector carry cables which are coupled with tension springs between a pair of rotatable booms. The tension springs cause the cables to assume a catenary shape which enhances the flatness of the reflector film. In order to fold the solar panels into a stowed position, the booms rotate to lie alongside the panels and the thin-film reflector is rolled (e.g., from the reflector center) so that it lies parallel to the booms. Although this reflector system is potentially lighter and simpler than the system described above, it still involves numerous mechanical parts (e.g., booms, cables and pulleys) which have significant weight.
Other reflector systems are described in U.S. patent application Ser. No. 08/081,909, filed Jun. 18, 1993, and now abandoned (as a continuation of application Ser. No. 07/802,972, filed Dec. 6, 1991 and now abandoned), titled "Augmented Solar Array with Dual Purpose Reflectors" and assigned to Hughes Electronics, the assignee of the present invention. In an exemplary system, a reflector is formed from a reflective material (e.g., an aluminized polyimide film) that is carried by a peripheral frame or affixed over a ribbed structure or a thin metal sheet. Each reflector is coupled to a solar panel by a hinge mechanism. Prior to spacecraft launch, the reflector is rotated to lie proximate to the solar cell face of the solar panel. After launch, the hinge mechanism rotates the reflector to a position in which it forms a deployment angle with the solar cell face. In an exemplary hinge mechanism, a hinge spring urges the reflector to rotate away from the solar cell face. The hinge mechanism includes a stop member which halts this rotation when the reflector reaches the deployment angle.
In another reflector system embodiment, reflectors are fabricated by suspending a reflective film between a pair of flexible rods that are rigidly coupled to a solar panel. The rods are typically tethered such that the reflectors lie parallel to the solar cell face prior to spacecraft launch. Deployment is effected by untethering which allows the rods to whip directly to a position in which the reflective film forms a deployed angle with the panel.
Although the latter reflector system effectively redirects radiation, the solar reflectors are stowed over the solar cell face of the solar panels. Accordingly, they block the use of the solar panels during any period (e.g., a transfer orbit) in which the solar panels are in a storage position that prevents reflector deployment.